The present invention relates to the use of generators to produce electrical power or thrust in aerospace craft, and more particularly, to generators for converting solar power to electrical power or thrust for a satellite.
Satellite systems typically have to operate a range of electrical devices and generate thrust in order to perform the various tasks laid out for them in a mission. For instance, a supply of propellant must be heated to produce the thrust necessary to move the satellite, or to change the attitude of the satellite. In addition, electrical devices used for various tasks require a supply of electricity. It is highly desirable to conserve satellite size and mass due to the need to place each satellite into orbit via an expensive launch vehicle. Therefore, the use of readily available solar power is typically desired over launching batteries, or other heavy and bulky power storage equipment. Prior art satellite systems have disclosed separate power and thrust generation systems which generally present a less efficient use of the limited size and mass payload capabilities of launch vehicles.
U.S. Pat. No. 5,459,996 to Malloy, III et al. discloses a satellite system that uses the same thermal storage device for alternatively generating thrust or electrical power. The satellite system includes a pair of mirror assemblies 50 that focus sunlight through a pair of windows 36 defined by a receiver 30. The receiver includes a hollow thermal storage drum 32 that defines an internal wall of the receiver, as shown in FIG. 2 of Malloy, III et al. The thermal storage drum absorbs radiant energy from sunlight entering through the windows, converts the radiant energy into thermal energy and subsequently retains the thermal energy for generation of electrical power or thrust. The thermal storage drum is encased by a removable diode insulation sleeve 40 which is retractable over the outer periphery of the thermal storage drum. An energy conversion medium 42 is located at the outer periphery of the insulation sleeve and converts the radiant energy stored in the thermal storage drum into electrical power when the insulation sleeve is retracted, as shown in FIG. 4 of Malloy, III et al.
During an alternative thrust mode, the insulation sleeve is returned to its original (unretracted) position, as shown in FIG. 3, which halts the production of electrical power. Propellant from a propellant supply is routed through the tubes 22 at the bottom of the receiver so that the propellant removes heat produced and stored by the thermal storage drum. The heated propellant is routed out of the receiver and into a nozzle 28 at the top of the receiver so as to produce thrust. Although advantageously reducing the mass of the satellite by combining thrust and electrical power generation in a single receiver, the satellite system disclosed by Malloy, III et al. does not generate thrust and electrical power simultaneously.
It would be advantageous to have a combined thrust and electrical power generation device for a satellite system. Further, it would be advantageous if the combined generation device were capable of simultaneously generating both thrust and electrical power for a satellite system. Such a simultaneous generation of thrust and electrical power would allow full operation of the satellite""s electrical devices even during movement of the satellite, or changes in its attitude. It would be further advantageous if the generation device used ambient solar radiation to generate the thrust and electrical power, reducing the need to carry additional fuel or power storage devices during launch of the satellite system into orbit.
The present invention addresses the above needs and achieves other advantages by providing a combined generator device for use in a satellite capable of generating electrical power and thrust from an optical solar image. The combined generator includes a first receiver positioned to receive and store the outer, low-intensity region of the optical solar image. An electrical power generator thermally coupled to the first receiver uses the stored thermal energy to produce electricity. The combined generator also includes a thrust generator having a second receiver. The second receiver is positioned to receive and absorb the inner, high intensity region of the optical solar image. A heat exchanger of the combined generator transfers the thermal energy from the second receiver into a supply of propellant which is vented to a nozzle to produce thrust.
In one embodiment, the present invention includes a combined generator device for use in a satellite. The combined generator device is capable of generating electrical power and propulsion from an optical solar image having a high-intensity region and a low-intensity region. A solar concentrator of the combined generator device is configured to direct and focus the two regions of the optical solar image onto a focal plane. Positioned at the focal plane is a first receiver that includes a first radiant energy absorbing member. The first radiant energy absorbing member is positioned so as to receive the focused low-intensity region of the optical solar image. The first receiver further includes a thermal storage medium operably connected to the first radiant energy absorbing member so as to be able to store the thermal energy absorbed by the first radiant energy absorbing member. An electrical power generator of the combined generator device is thermally coupled to the first receiver, receives thermal energy from the first receiver, and is configured to convert the thermal energy into electrical power. The combined generator device also includes a propulsion generator having a second receiver and a propellant supply. The second receiver includes a second radiant energy absorbing member positioned at the focal plane so as to receive the high-intensity region of the optical solar image. The propellant supply is thermally coupled with the second receiver and receives thermal energy from the second receiver which heats the propellant to produce thrust.
The first radiant energy absorbing member may define a first cavity having an aperture positioned at the focal plane. In such a case, the second radiant energy absorbing member of the second receiver is positioned within the first cavity. Preferably, the first cavity defined by the first radiant energy absorbing member is a cylindrical cavity, wherein the second receiver is concentrically positioned within the first cavity. The second radiant energy absorbing member may also define a cylindrical cavity. The thermal energy storage medium, such as graphite, is layered around an outer surface of the first radiant energy absorbing member. The first and second radiant energy members are constructed of a refractory material, such as rhenium.
The thrust generator may include a heat exchanger thermally coupled with the second radiant energy absorbing member and in fluid communication with the propellant supply. Propellant passing through the heat exchanger is heated by the second radiant energy absorbing member so as to produce thrust. To aid in the production of thrust, the propellant may be pre-heated by a pre-heat exchanger thermally coupled with the first radiant energy absorbing member. The pre-heat exchanger is also connected in fluid communication with the propellant supply, but upstream of the heat exchanger. After heating, the propellant is typically directed by a thrust nozzle that is in fluid communication with the propellant supply and is downstream from both the pre-heat exchanger and the heat exchanger.
The present invention has several advantages. The combined thrust and power generator can simultaneously generate thrust and electrical power by using the high-intensity and low-intensity regions of the solar image. Such use of the naturally varying intensity of the ambient solar image optimizes the efficient use of radiant heat generated by the solar image. In other words, the low-intensity region of the solar image is not wasted, but is directed to electrical power generation which does not require the thermal energy associated with the higher temperature, while the high intensity region of the solar energy is directed to thrust generation which prefers the thermal energy associated with the higher temperature. In addition, the combination of functions reduces the size and mass of the generator, which reduces the expense of launching the satellite. No moving parts are required to alternate between two different power and thrust generation modes, increasing the reliability of the combined generator. A circular first and second radiant energy absorbing members of one advantageous embodiment match the circular solar image, thereby reducing any loss of the solar image that fails to hit an energy absorbing member. Pre-heating of the propellant reduces the thermal energy needed to achieve a target propellant temperature for the production of thrust.